Gas turbine engine with noise attenuating variable area fan nozzle

ABSTRACT

A nacelle assembly for a high-bypass gas turbine engine includes a core nacelle defined about an engine centerline axis. A fan nacelle is mounted at least partially around the core nacelle to define a fan bypass flow path. A variable area fan nozzle is in communication with the fan bypass flow path. The variable area fan nozzle has a first fan nacelle section and a second fan nacelle section. The second fan nacelle section is axially movable relative to the first fan nacelle section to define an auxiliary port at a non-closed position to vary a fan nozzle exit area and adjust fan bypass airflow. The second fan nacelle section includes an acoustic system that has an acoustic impedance located on a radially outer surface.

CROSS-REFERENCE TO RELATED APPLICATIONS

This disclosure is a continuation of U.S. patent application Ser. No.12/147,432 filed Jun. 26, 2008.

BACKGROUND OF THE INVENTION

The present invention relates to a gas turbine engine, and moreparticularly to a turbofan engine having a variable area fan nozzle(VAFN) with an acoustic system to attenuate leading edge noise andreduce the total effective perceived noise level (EPNL).

Gas turbine engines which have an engine cycle modulated with a variablearea fan nozzle (VAFN) provide a smaller fan exit nozzle diameter duringcruise conditions and a larger fan exit nozzle diameter during take-offand landing conditions.

The VAFN may generate significant noise as upstream turbulence interactswith the leading edge of the VAFN. The upstream turbulence may resultfrom turbulent boundary layers which expand from the upstream fixednacelle wall, turbulence which evolves from the upstream fan exit guidevane (FEGV) wakes or endwall effects, and flow separation that may occurfrom the contour of the upstream nacelle wall. The physical mechanismfor leading edge VAFN noise, which exhibits acoustic dipole behavior, isfundamentally different from traditional jet exhaust mixing noise, whichexhibits acoustic quadrupole behavior. Additionally, this excess noiseis not significantly reduced in forward flight as typical jet exhaustmixing noise. Thus, the leading edge source is of significant importancein its contribution toward the effective perceived noise level (EPNL).

SUMMARY OF THE INVENTION

In one exemplary embodiment, a nacelle assembly for a high-bypass gasturbine engine includes a core nacelle defined about an enginecenterline axis. A fan nacelle is mounted at least partially around thecore nacelle to define a fan bypass flow path. A variable area fannozzle is in communication with the fan bypass flow path. The variablearea fan nozzle has a first fan nacelle section and a second fan nacellesection. The second fan nacelle section is axially movable relative tothe first fan nacelle section to define an auxiliary port at anon-closed position to vary a fan nozzle exit area and adjust fan bypassairflow. The second fan nacelle section includes an acoustic system thathas an acoustic impedance located on a radially outer surface.

In a further embodiment of the above, the acoustic system is defined atleast in part within a leading edge region of the second fan nacellesection.

In a further embodiment of any of the above, the acoustic system furthercomprises a forward acoustic system and an aft acoustic system. Theforward acoustic system is different than the aft acoustic system.

In a further embodiment of any of the above, the acoustic system furthercomprises a forward acoustic system and an aft acoustic system. Theforward acoustic system comprises the leading edge of the second fannacelle section. The aft acoustic system comprises at least a portion ofan upper surface of the second fan nacelle section.

In a further embodiment of any of the above, the acoustic systemincludes a perforated inner face sheet and a perforated outer face sheetsupported by a structure.

In a further embodiment of any of the above, the acoustic systemcomprises a perforated inner face sheet.

In a further embodiment of any of the above, the acoustic systemcomprises an outer face sheet.

In a further embodiment of any of the above, the acoustic systemcomprises a bulk absorbing material.

In a further embodiment of any of the above, the acoustic systemcomprises a forward acoustic system that is in fluid communication withan aft acoustic system. The forward acoustic system comprises a bulkabsorbing material and said aft acoustic system comprises a perforatedouter face sheet along at least a portion of an upper surface of saidsecond fan nacelle section.

In a further embodiment of any of the above, the second fan nacellesection defines a trailing edge of the variable area fan nozzle.

In a further embodiment of any of the above, the second fan nacellesection is subdivided into a multiple of independently operable sectors.Each of the multiple of independently operable sectors is axiallymovable relative to the first fan nacelle section to define anasymmetric fan nozzle exit area.

In another exemplary embodiment, a high-bypass gas turbine engineincludes a core engine defined about an axis. A gear system is driven bythe core engine. A turbofan is driven by the gear system about the axis.A core nacelle is defined at least partially about the core engine. Afan nacelle is mounted at least partially around the core nacelle todefine a fan bypass flow path. A variable area fan nozzle is incommunication with the fan bypass flow path. The variable area fannozzle has a first fan nacelle section and a second fan nacelle section.The second fan nacelle section is axially movable relative to the firstfan nacelle section to define an auxiliary port at a non-closed positionto vary a fan nozzle exit area and adjust fan bypass airflow. The secondfan nacelle section includes an acoustic system that is located on aleading edge and radially outer surface.

In a further embodiment of any of the above, the acoustic systemcomprises a perforated inner face sheet and a perforated outer facesheet that is supported by a structure.

In a further embodiment of any of the above, the acoustic systemcomprises a bulk absorbing material.

In a further embodiment of any of the above, the second fan nacellesection defines a trailing edge of the variable area fan nozzle.

In another exemplary embodiment, a method of reducing a total effectiveperceived noise level of a gas turbine engine with a variable area fannozzle. The method includes axially moving a second fan nacelle sectionbetween a closed position in which the second fan nacelle section is insequential alignment with a first fan nacelle section in response to acruise flight condition and an open position in which the second fannacelle section is aftward of the first fan nacelle section to define anauxiliary port. The second fan nacelle section has a leading edge regionwith an acoustic system which provides an acoustic impedance when thesecond fan nacelle section is positioned at a non-closed position.

In a further embodiment of any of the above, the method includesgenerating the acoustic impedance with at least a portion of an uppersurface of the second fan nacelle section.

In a further embodiment of any of the above, the open position is inwhich the second fan nacelle section is aftward of the first fan nacellesection to define an auxiliary port is in response to a non-cruiseflight condition.

BRIEF DESCRIPTION OF THE DRAWINGS

The various features and advantages of this invention will becomeapparent to those skilled in the art from the following detaileddescription of the currently preferred embodiment. The drawings thataccompany the detailed description can be briefly described as follows:

FIG. 1A is a general schematic partial fragmentary view of an exemplarygas turbine engine embodiment for use with the present invention;

FIG. 1B is a rear view of the engine;

FIG. 1C is a side view of the engine integrated with a pylon;

FIG. 1D is a rear perspective view of the engine integrated with apylon;

FIG. 2A is a sectional side view of the VAFN in a closed position;

FIG. 2B is a sectional side view of the VAFN in an open position; and

FIG. 3 is a sectional side view of the VAFN with an acoustic system;

FIG. 4 is a sectional side view of one non-limiting embodiment of theacoustic system;

FIG. 5 is a sectional side view of another non-limiting embodiment ofthe acoustic system; and

FIG. 6 is a sectional side view of yet another non-limiting embodimentof the acoustic system.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT

FIG. 1A illustrates a general partial fragmentary schematic view of agas turbofan engine 10 suspended from an engine pylon P within an enginenacelle assembly N as is typical of an aircraft designed for subsonicoperation.

The turbofan engine 10 includes a core engine within a core nacelle 12that houses a low spool 14 and high spool 24. The low spool 14 includesa low pressure compressor 16 and low pressure turbine 18. The low spool14 also drives a fan section 20 through a gear train 22. The high spool24 includes a high pressure compressor 26 and high pressure turbine 28.A combustor 30 is arranged between the high pressure compressor 26 andhigh pressure turbine 28. The low and high spools 14, 24 rotate about anengine axis of rotation A.

The engine 10 in one non-limiting embodiment is a high-bypass gearedarchitecture aircraft engine with a bypass ratio greater than ten(10:1), a turbofan diameter significantly larger than that of the lowpressure compressor 16, and the low pressure turbine 18 with a pressureratio greater than 5:1. The gear train 22 may be an epicycle gear trainsuch as a planetary gear system or other gear system with a gearreduction ratio of greater than 2.5:1. It should be understood, however,that the above parameters are only exemplary of one non-limitingembodiment of a geared architecture engine and that this disclosure isapplicable to other gas turbine engines including direct driveturbofans.

Airflow enters a fan nacelle 34, which at least partially surrounds thecore nacelle 12. The fan section 20 communicates airflow into the corenacelle 12 to power the low pressure compressor 16 and the high pressurecompressor 26. Core airflow compressed by the low pressure compressor 16and the high pressure compressor 26 is mixed with the fuel in thecombustor 30 and expanded over the high pressure turbine 28 and lowpressure turbine 18. The turbines 28, 18 are coupled for rotation with,respective, spools 24, 14 to rotationally drive the compressors 26, 16and through the gear train 22, the fan section 20 in response to theexpansion. A core engine exhaust E exits the core nacelle 12 through acore nozzle 43 defined between the core nacelle 12 and a tail cone 32.

The core nacelle 12 is supported within the fan nacelle 34 bycircumferentially space structures 36 often generically referred to asFan Exit Guide Vanes (FEGVs). A bypass flow path 40 is defined betweenthe core nacelle 12 and the fan nacelle 34. The engine 10 generates ahigh bypass flow arrangement with a bypass ratio in which approximatelyeighty percent of the airflow which enters the fan nacelle 34 becomesbypass flow B. The bypass flow B communicates through the generallyannular bypass flow path 40 and is discharged from the engine 10 througha variable area fan nozzle (VAFN) 42 which defines a nozzle exit area 44between the fan nacelle 34 and the core nacelle 12 at a fan nacelle endsegment 34S of the fan nacelle 34 downstream of the fan section 20.

Thrust is a function of density, velocity, and area. One or more ofthese parameters can be manipulated to vary the amount and direction ofthrust provided by the bypass flow B. The VAFN 42 operates toeffectively vary the area of the fan nozzle exit area 44 to selectivelyadjust the pressure ratio of the bypass flow B in response to acontroller C. Low pressure ratio turbofans are desirable for their highpropulsive efficiency. However, low pressure ratio fans may beinherently susceptible to fan stability/flutter problems at low powerand low flight speeds. The VAFN allows the engine to change to a morefavorable fan operating line at low power, avoiding the instabilityregion, and still provide the relatively smaller nozzle area necessaryto obtain a high-efficiency fan operating line at cruise.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 20 of the engine 10 is preferablydesigned for a particular flight condition—typically cruise at 0.8M and35,000 feet. As the fan blades within the fan section 20 are efficientlydesigned at a particular fixed stagger angle for an efficient cruisecondition, the VAFN 42 is operated to effectively vary the fan nozzleexit area 44 to adjust fan bypass air flow such that the angle of attackor incidence on the fan blades is maintained close to the designincidence for efficient engine operation at other flight conditions,such as landing and takeoff to thus provide optimized engine operationover a range of flight conditions with respect to performance and otheroperational parameters such as noise levels.

The VAFN 42 may be separated into at least two sectors 42A-42B (FIG. 1B)defined between the pylon P and a lower Bi-Fi splitter L which typicallyinterconnects a larger diameter fan duct reverser cowl and a smallerdiameter core cowl (FIGS. 1C and 1D). Each of the at least two sectors42A-42B are independently adjustable to asymmetrically vary the fannozzle exit area 44 to generate vectored thrust. It should be understoodthat although two segments are illustrated, any number of sectors andsegments may alternatively or additionally be provided.

The VAFN 42 generally includes an auxiliary port system 50 having afirst fan nacelle section 52 and a second fan nacelle section 54 movablymounted relative the first fan nacelle section 52. The second fannacelle section 54 axially slides along the engine axis A relative thefixed first fan nacelle section 52 to change the effective area of thefan nozzle exit area 44. The second fan nacelle section 54, in onenon-limiting embodiment, slides aftward upon a track fairing 56A, 56B(illustrated schematically in FIGS. 1C and 1D) in response to anactuator 58 (illustrated schematically). The track fairing 56A, 56Bextend from the first fan nacelle section 52 adjacent the respectivepylon P and the lower Bi-Fi splitter L (FIG. 1D).

The VAFN 42 changes the physical area and geometry of the bypass flowpath 40 during particular flight conditions. The bypass flow B iseffectively altered by sliding of the second fan nacelle section 54relative the first fan nacelle section 52 between a closed position(FIG. 2A) and an open position (FIG. 2B). The auxiliary port system 50is closed by positioning the second fan nacelle section 54 in-line withthe first fan nacelle section 52 to define the fan nozzle exit area 44as exit area F0 (FIG. 2A).

The VAFN 42 is opened by moving the second fan nacelle section 54aftward along the track fairing 56A, 56B away from the first fan nacellesection 52 to open an auxiliary port 60 (FIG. 2B) which extends betweenthe open second fan nacelle section 54 relative the first fan nacellesection 52 to essentially provide an increased fan nozzle exit area 44exit area F1. That is, the exit area F1 with the auxiliary port 60 (FIG.2B) is greater than exit area F0 (FIG. 2A).

In one non-limiting embodiment, the auxiliary port 60 is incorporatedwithin the bypass flow path 40 aft of the Fan Exit Guide Vanes 36(FEGVs). The auxiliary port 60 is located through the bypass duct outerwall.

In operation, the VAFN 42 communicates with the controller C to move thesecond fan nacelle section 54 relative the first fan nacelle section 52of the auxiliary port system 50 to effectively vary the area defined bythe fan nozzle exit area 44. Various control systems including an enginecontroller or an aircraft flight control system may also be usable withthe present invention. By adjusting the axial position of the entireperiphery of the second fan nacelle section 54 in which all sectors aremoved simultaneously, engine thrust and fuel economy are maximizedduring each flight regime by varying the fan nozzle exit area. Byseparately adjusting the sectors of the second fan nacelle section 54 toprovide an asymmetrical fan nozzle exit area 44, engine bypass flow isselectively vectored to provide, for example only, trim balance, thrustcontrolled maneuvering, enhanced ground operations and short fieldperformance.

Referring to FIG. 3, the second fan nacelle section 54 includes aleading edge region 62 with an acoustic system 64. The acoustic system64 utilizes the available volume of the leading edge region 62 toachieve an optimal acoustic impedance. It should be understood that themaximum pressure difference across the VAFN primarily occurs in theleading or forward one-third of the second fan nacelle section 54 andthat the leading edge region 62 includes at least that area. Withoptimal acoustic impedance, the acoustic system 64 operates to avoidsource radiation from the leading edge and/or attenuate the leading edgenoise which thus avoids propagation to the far-field which reduces thetotal effective perceived noise level (EPNL).

Referring to FIG. 4, one non-limiting embodiment of the acoustic system64A includes a perforated inner face sheet 66 and a perforated outerface sheet 68 supported by a structure 70. The micro-porosity of theperforated inner face sheet 66, the micro-porosity of the perforatedouter face sheet 68 and the arrangement of the structure 70 are arrangedto tune the acoustic system 64A to provide an optimal acoustic impedanceand achieve maximum attenuation. The structure 70 ensures local reactioncharacteristics within the leading edge region 62 through acousticcommunication between one or both of the perforated inner face sheet 66and the perforated outer face sheet 68. It should be understood that thestructure 70 is illustrated in partial schematic cross-section and thatvarious arrangements of the structure 70 may be provided to support theperforated inner face sheet 66 and the perforated outer face sheet 68.Although both the perforated inner face sheet 66 and the perforatedouter face sheet 68 are illustrated in a single non-limiting embodiment,it should be understood that only one or both of the perforated innerface sheet 66 and the perforated outer face sheet 68 may be utilized.

Referring to FIG. 5, another non-limiting embodiment of the acousticsystem 64B includes a bulk absorbing material 80 such as, for exampleonly, a sintered metal, a ceramic foam, Kevlar, or a carbide material tominimize effects on the steady flow through the auxiliary port 60 andmaximize effects on unsteady loading. As with the FIG. 4 embodiment, theporosity, depth, and material characteristics are selected for optimalimpedance and thus optimal acoustic attenuation.

Referring to FIG. 6, another non-limiting embodiment of the acousticsystem 64C includes a forward acoustic system 90 and an aft acousticsystem 92. The aft acoustic system 92 may provide an additional surfacearea to supplement performance of the forward acoustic system 90. Theforward acoustic system 90 may be the same or different from the aftacoustic system 92. That is, the forward acoustic system 90 may beeither the FIG. 4 perforated plate design or the bulk absorber design ofFIG. 5 coupled with the aft acoustic system 92 which may be either theFIG. 4 perforated plate design or the bulk absorber design of FIG. 5.

In this non-limiting embodiment, the aft acoustic system 92 is locatedonly along an outer surface 54A of the second fan nacelle section 54.The aft acoustic system 92 may alternatively or additionally include aperforated plate design bounded to partitions in the internal volume ora wire mesh acoustic liner, if a relatively larger attenuation bandwidthis desired.

Noise reduction on the order of approximately 3 EPNdB cumulative overthe certification conditions described in Federal Acquisition Regulation(FAR) 36 may be readily achieved by the acoustic system 64 disclosedherein and include both tone and broadband reductions.

The foregoing description is exemplary rather than defined by thelimitations within. Many modifications and variations of the presentinvention are possible in light of the above teachings. The preferredembodiments of this invention have been disclosed, however, one ofordinary skill in the art would recognize that certain modificationswould come within the scope of this invention. It is, therefore, to beunderstood that within the scope of the appended claims, the inventionmay be practiced otherwise than as specifically described. For thatreason the following claims should be studied to determine the truescope and content of this invention.

What is claimed is:
 1. A nacelle assembly for a high-bypass gas turbineengine comprising: a core nacelle defined about an engine centerlineaxis; a fan nacelle mounted at least partially around said core nacelleto define a fan bypass flow path; and a variable area fan nozzle incommunication with said fan bypass flow path, said variable area fannozzle having a first fan nacelle section and a second fan nacellesection, said second fan nacelle section axially movable relative saidfirst fan nacelle section to define an auxiliary port at a non-closedposition to vary a fan nozzle exit area and adjust fan bypass airflow,said second fan nacelle section including an acoustic system having anacoustic impedance located on a radially outer surface, wherein theacoustic system is located on a fore portion of the radially outersurface, an aft portion of the radially outer surface lacking theacoustic system.
 2. The nacelle assembly as recited in claim 1, whereinsaid acoustic system is defined at least in part within a leading edgeregion of said second fan nacelle section.
 3. The nacelle assembly asrecited in claim 1, wherein said acoustic system further comprises aforward acoustic system and an aft acoustic system, said forwardacoustic system being different than said aft acoustic system.
 4. Thenacelle assembly as recited in claim 1, wherein said acoustic systemfurther comprises a forward acoustic system and an aft acoustic system,said forward acoustic system comprises said leading edge of said secondfan nacelle section and said aft acoustic system comprises at least aportion of an upper surface of said second fan nacelle section.
 5. Thenacelle assembly as recited in claim 1, wherein said acoustic systemcomprises a perforated inner face sheet and a perforated outer facesheet supported by a structure.
 6. The nacelle assembly as recited inclaim 1, wherein said acoustic system comprises a perforated inner facesheet.
 7. The nacelle assembly as recited in claim 1, wherein saidacoustic system comprises an outer face sheet.
 8. The nacelle assemblyas recited in claim 1, wherein said acoustic system comprises a bulkabsorbing material.
 9. A nacelle assembly for a high-bypass gas turbineengine comprising: a core nacelle defined about an engine centerlineaxis; a fan nacelle mounted at least partially around said core nacelleto define a fan bypass flow path; and a variable area fan nozzle incommunication with said fan bypass flow path, said variable area fannozzle having a first fan nacelle section and a second fan nacellesection, said second fan nacelle section axially movable relative saidfirst fan nacelle section to define an auxiliary port at a non-closedposition to vary a fan nozzle exit area and adjust fan bypass airflow,said second fan nacelle section including an acoustic system having anacoustic impedance located on a radially outer surface; wherein saidacoustic system comprises a forward acoustic system in fluidcommunication with an aft acoustic system, said forward acoustic systemcomprises a bulk absorbing material and said aft acoustic systemcomprises a perforated outer face sheet along at least a portion of anupper surface of said second fan nacelle section.
 10. The nacelleassembly as recited in claim 9, wherein said second fan nacelle sectiondefines a trailing edge of said variable area fan nozzle.
 11. Thenacelle assembly as recited in claim 9, wherein said second fan nacellesection is subdivided into a multiple of independently operable sectors,each of said multiple of independently operable sectors axially movablerelative to said first fan nacelle section to define an asymmetric fannozzle exit area.
 12. A high-bypass gas turbine engine comprising: acore engine defined about an axis; a gear system driven by said coreengine; a turbofan driven by said gear system about said axis; a corenacelle defined at least partially about said core engine; a fan nacellemounted at least partially around said core nacelle to define a fanbypass flow path; and a variable area fan nozzle in communication withsaid fan bypass flow path, said variable area fan nozzle having a firstfan nacelle section and a second fan nacelle section, said second fannacelle section axially movable relative said first fan nacelle sectionto define an auxiliary port at a non-closed position to vary a fannozzle exit area and adjust fan bypass airflow, said second fan nacellesection including an acoustic system located on a leading edge andradially outer surface, wherein the acoustic system is located on a foreportion of the radially outer surface, an aft portion of the radiallyouter surface lacking the acoustic system.
 13. The high-bypass gasturbine engine as recited in claim 12, wherein said acoustic systemcomprises a perforated inner face sheet and a perforated outer facesheet supported by a structure.
 14. The high-bypass gas turbine engineas recited in claim 12, wherein said acoustic system comprises a bulkabsorbing material.
 15. The high-bypass gas turbine engine as recited inclaim 12, wherein said second fan nacelle section defines a trailingedge of said variable area fan nozzle.
 16. A method of reducing a totaleffective perceived noise level of a gas turbine engine with a variablearea fan nozzle comprising: axially moving a second fan nacelle sectionbetween a closed position in which said second fan nacelle section is insequential alignment with a first fan nacelle section in response to acruise flight condition and an open position in which said second fannacelle section is aftward of said first fan nacelle section to definean auxiliary port, said second fan nacelle section having an acousticsystem located on a radially outer surface which provides an acousticimpedance when said second fan nacelle section is positioned at anon-closed position, wherein the acoustic system is located on a foreportion of the radially outer surface, an aft portion of the radiallyouter surface lacking the acoustic system.
 17. The method as recited inclaim 16, further comprising: generating said acoustic impedance with atleast a portion of an upper surface of said second fan nacelle section.18. The method as recited in claim 16, wherein said open position inwhich said second fan nacelle section is aftward of said first fannacelle section to define an auxiliary port is in response to anon-cruise flight condition.